Aircraft with autorotative wings



y 1945. J. DE LA ClERVA v AIRCRAFT WITH AUTOROTATIVE WINGS Filled Aug. 5, 1934 i 6 Shets-Sheet 1 4150M ATTORNEYJ July 31, 1945. J. DE LA ClERVA AIRCRAFT WITH AUTOROTATIVE WINGS Filed Aug. 3, 1934 6 Sheets-Sheet 2 INVENTOR. MW

ATTORNEY) July 31, 1945. J. DE LA.C XERVA 2,380,583

AIRCRAFT WITH AUTOROTATSIIVE WI NGS Filed Aug. 5; 1934 I 6 Sheets-Sheet 3 l N VEN TOR.

W ge 45m ATTORNEY;

, July 1945' Y J. DE LA CIERVA 2,380,583

AIRCRAFT WITH AUTOROTATIVE WINGS Filed Aug. 5, 1954 I s Sheets-Sheet 4 f I i i I i I 1 (Y &

,%I% ,/.fi i I /"i R, '1" I E R 0 Q Q Q IN VEN TOR.

j) A ammy;

July 1945' J. DE LA CIERVA 2,380,583

AIRCRAFT WITH AUTO ROT AT IVE WINGS Filed Aug. s,'1954 6 Sheets-Sheet 5 IIIVVENTOR.

M BY/ ATTORNEYJ M 31, 1945- J. DE LA CIERVA AIRCRAFT WITH AUTOROTATIVE WINGS Filed Aug. 5, 1934 6 Sheets-Sheet 6 INVENTOR.

ATTORNEYJ.

Patented July 31, 1945 2.380.583 smcasr'r wrrn AUTOROTA'HVE wmGs Juan de la Cierva, Madrid, Spain, assignor, by

memo assignments, to Autogiro Company America, a corporation of Delaware Application August 3, 1934, Serial No. 738,349

- In Great Britain Anmt 5, 1933 so Clalmsr" (01.244-18) The present" invention relates to aircraft in which the principal meanspf support in flight consists of a system of rotatlve wings or, blades,

hereinafter referred toas a rotor, mounted for a disconnectible driving connection between the said engine and the rotor adapted to apply a torque to said rotor.

More particularly this invention refers to aircraft having sustaining rotors of the kind referred to in which the individual blades are attached to the hub or central member of the rotor by flexible or articular connections which permit each blade to swing up and down substantially in a plane containing the rotor axis.

The principal object of the invention is to improve the take-off of an aircraft of the kind above referred to by eliminating the run along the ground hitherto necessary to enable the rotor to develop enough lift to raise the aircraft after being rotated by the starting means; in other words to provide this type of craft with the capability of vertical take-oil as well as vertical descent.

According to the present invention, in an aircraft having a sustaining rotor of the kind referred to and power transmission means for applying a driving torque to the rotor for the purpose of imparting thereto an initial rotation prior to flight, the maximum torque so transmissible to the rotor being less than that required for helicopter flight (i. e. sustained hovering), the power transmission means are adapted to impart to the rotor an initial rotational speed considerably in excess of the mean autorotativ'e speed in flight and means are provided for substantially decreasing the rotational drag of the .rotor during the application of the starting torque and for utilizing the excess kinetic energy stored in the rotor by virtue of its initial excess of rotational speed to generate temporarily a thrust sufliclent to raise the aircraft vertically from the ground by changing the mean pitch angle of the rotor blades so that during the application of the starting torque the said mean pitch angle may be decreased to a value substantially less than that required in flight and on the van hiring of said. starting torque said mean pitch l8 angle may be immediately increased at least to the minimum value required for flight.

According to a feature of the invention the means for changing the pitch angle of the rotor blades operate automatically in response to variatfons of the torque applied to the rotor.

Preferably the application of a substantial driving torque to the rotor, at least with the rotor turning at less than the minimum autorotatlonal speed of flight, brings the mean pitch angle of the rotor blades substantially to the zero value, i. e. the value for minimum aerodynamic rotational drag.

According to another feature of the invention the means for ehangingthe pitch angle of the rotor blades operate automatically in response to changes of the centrifugal force experienced by the blades, an increase of centrifugal force causing the mean pitch angle to increase and vice versa.

If the automatic rotor blade pitch changing means are responsive to changes both of torque and of centrifugal force the construct ion preferably ensures that the influence of the torque transmitted to the rotor so preponderates over the influence of centrifugal force that the application of maximum torque sufllces to maintain the mean blade pitch angle at or near its minimum value at least while the rotational speed of the rotor does not exceed the minimum autorotational speed of flight.

By reducing the pitch angle of the rotor blades substantially to that of zero lift during thev application of the starting torque (that is, zero lift position with respect to a plane perpendicular to the axis of rotation of the rotor), the aerodynamic resistance of the blades is reduced to its minimumv value whereby the required increased speed of initial rotation can be imparted to the rotor without increased expenditure of power.

Since. with a moderate expenditure of power, the rotor can be given an initial speed of rotation much in excess of the normal rotational speed in flight, if the pitch angle of the blades is also influenced by the centrifugal force, in the manner above referred to, the excess centrifugal force (or kinetic energy of rotation) will on the disappearance of the starting torque cause the blades to assume a pitch angle greater than the normal in flight, whereby an increased lifting force is obtained. The pitch angle will of course revert to the normal as soon as the excess energy of the rotor has been expended and its speed of rotation has reverted to its normal autorotativc speed in flight.

Further, the control of the pitch angles of the blades in accordance with the centrifugal force as above referred to, confers a further advantage in that it causes the pitch angles of the blades to adjust themselves more nearly to the optimum values for all conditions of normal flight. Thus, it is known that with an autorotative rotor of the kind herein first referred to, the speed of rotation increases somewhat with increase of forward flight speed of the aircraft. With the arrangement of the present invention this increase of rotational speed with forward speed, by reason of the consequent increase of centrifugal force, causes the pitch angles of the rotor blades to increase and it is known both from theory and experiments that the optimum pitch angle for aerodynamic efficiency of an autorotative rotor varies with the speed of translation, being greater the higher the translational speed. Thus by suitably choosing the ratio which increase of pitch angle bears to increase of centrifugal force the rotor blades may be made to assume the optimum pitch angle for aerodynamic efficiency throughout substantially the whole of the translationa speed range. I

Another advantage that follows from the control of pitch angle in accordance with centrifugal force as above described, consists in that any cause tending to decrease the rotational speed of the blades is immediately compensated by a decrease in pitch angle resulting from the deficiency of centrifugal force. Thus the danger of stoppage or excessive slowing of the rotor following any accidental cause tending to increase the drag of the rotor blades, as for instance damage to a blade or the formation of ice on the blades, is largely obviated.

Means are preferably provided setting an upper limit for the variation of mean pitch angle of the rotor blades, which upper limit does not exceed the upper limiting value of the range of rotor blade pitch angles within which continuous autorotation of the rotor is possible in steady, i. e. unaccelerated, flight conditions. The upper limit of rotor blade pitch angle (with the usualv aerofoil sections employed at present commercially in autorotative-winged machines) should not exceed eight degrees, measured from the pitch angle of zero lift coeflicient.

According to a further feature of the invention restraining means are provided, operative.

during the application of the starting torque, for preventing the pitch angle of the rotor blades from increasing to a value sufilcient to raise the aircraft from the ground as a result of an accidental momentary or transient diminution of the applied torque. The said means may operate automatically by applying a frictional resistance to change of rotor blade pitch angle at least when said pitch angle is at or near its minimum value, or may comprise a locking device operative automatically tolock the rotor,

asaqsss dividual rotor blades are articulated to the rotor hub member by pivotal means permitting independent swinging of each blade relatively to the hub member about two independent pivots of which one (hereinafter called the flapping" pivot) has its axis substantially in a plane perpendicular to the rotational axis and the other thereinaftencalled the "drag" pivot) has its axis substantially in a plane containing the rotational axis. Means are provided for varying the pitch angle of each rotor blade responsively to its movement on said drag pivot in such a way that movement of the blade relatively to the rotor hub in a direction contrary to the' direction of rotation of the rotor causes the blade pitch angle to decrease and vice versa.

It will be seen that when the driving torque is aplied to the hub the blades will tend to lag behind the hub in its direction of rotation, and this will have the effect of decreasing the pitch angle, thus reducing the aerodynamic resistance to rotation.

When the rotor is rotating under the influence of an applied driving torque the rotor blades will acquire a position of equilibrium in which the applied and resisting torques are balanced by the centrifugal force (it being understood that the drag pivot axes are offset from the rotational axis) and in this position each blade will have swung back from its mean position with respect to the hub through an angle hereinafter referred to as the angle of lag which is determined by the relative values of the applied torque and of the centrifugal force.

It will be noted that at the beginning of the application of the starting torque, the centrifugal force being small, the angle of lag will be large if the blade is free (unlocked) and since it is desirable that when the rotor has acquired its full starting revolutions the pitch angle of the blades shall not be substantially different from zero, stop means are preferably included to prevent the blades from taking up appreciable negative pitch angles at the beginning of the application of the starting torque. These stop means may either be arranged to limit the angle of lag or to limit the possible variation of pitch angle.

In fixing the position of these stop means allowance must be made for the possible tortional deflection of the rotor blades under load, which may, with the kind of blade having a forwardly disposed mass centre as hereinafter described, be such that the best setting of the stop'means gives a minimum blade pitch angle of two degreesat no load.

Any convenient mechanical devices may be employed for bringing about the variation of blade pitch angle in conformity with the variations in the angle of lag; such devices may include any suitable arrangements of linkages, cams and the like.

Whatever the mechanism employed it is desirable that the blades be free to oscillate about their drag pivots in flight under the action of aerodynamic and inertia. forces and that any limiting or stop means employed should not operate to interfere with such natural free oscil- IatiOIlS.

In a preferred form of construction, the desired interrelation of. pitch angle and angleof lag is obtained in a very simple manner by so constructing the blade and its articulations to the hub that the longitudinal axis of the blade is inclined upwards toward the axis of the secondary or drag pivot, so that the upper angle included between the said longitudinal axis and the said pivot axis is less than 90. Stated in another way, and viewing the blade as extending in a true horizontal radial position from the vertical hub axis (in other words, in a neutral position on the ordinary horizontal or flapping pivot, sometimes called the primary articulation) the drag pivot is extended substantially in a vertical plane containing the rotor axis, and the upward projection of the axis of said drag pivot is inclined outwardly and downwardly toward the blade axis so as to form an acute angle therewith.

This arrangement not only ensures an automatic decrease of pitch angle as. the blade lags behind the hub about its drag pivot and correspondingly an increase of pitch angle if the blade advances in front of the hub (or leads, in the direction of rotation), but also contributes toward the desired effect of interdependence of pitch angle and centrifugal force on the blade as above described.

The invention'also has reference to the construction of the rotor blades. In'the improved rotor blade construction according to this inabout the said pivot axis which causes the blade to lag. However, when the blade lags behind its mean position (in which the longitudinal blade axis cuts the rotational axis of the rotor) a centrifugal couple is created opposing the said lag. Theblade :will therefore take up or will oscillate about a position of equilibrium which is lagging. It is then evident that an increase of centrifugal force by decreasing the angle of lag in the equilibrium position will automatically bring about an increase of pitch angle.

It will further be seen that with this arrangement variations of lift on the blade tend to be neutralized since an increase of lift will, owing tothe inclination of the drag pivot, cause an increase of the lagging angleand consequently a diminution in pitch and vice versa. In this way variations of lift bring about pitch variations having the effect of compensating the lift variations. A reduction of flapping about the horizontal axis and a more uniform coning angle may thus be obtained, with a resultant increase of rotor efliciency.

With the object of improving as much as possible the take-off of the aircraft it is desirable to impart to the rotor as high an initial speed of rotation as possible with the available power vention the mass distribution and aerodynamic characteristics are such that the blade in flight tends to take up a position of equilibrium in which the centrifugal force exerts a couple about the drag pivot acting forwardly, i. e, in the direction of rotation, so that an increase of rotational speed causes the mean blade pitch angle to increase and vice versa. This may be accomplished by locating the mass center of v the blade in advance of the center of pressure;

and to ensure this under all conditions, preferably the mass center of the rotor blade is situated in advance of (i. e. nearer the leading edge than) the line with reference to which the pitching moment coefficient of the blade is substantially constant. In the aerofoil sections in general use for rotor blades for aircraft of the kind herein referred to this line lies at approximately 25% of the chord from the leading edge.- By placing the mass center forwardly of this line a pitching couple on the blade tending to decrease the pitch angle is assured at all angles of pitch with the rotor under load.

Thus when a rotor blade incorporates this feature and is mounted as above described with its drag pivot axis outwardly inclined down toward the blades longitudinal axis the pitch angle of the blade will increase with increase of centrifugal force and vice versa. The reason for this is as follows:

The offsetting of the aerodynamic center of pressure behind the mass center'of the blade introduces a pitching couple about the longitudinal which can be applied to the starter.

If the engine of the aircraft is coupled to a fixed pitch airscrew it is found that the maximum transmission of power to the rotor is obtained when the engine is running at about half its normal ground revolutions, i. e. the revolutions at which the air-screw alone absorbs the whole of the engine power when working at zero advance. (This conclusion of course depends on the shape of the engine full throttle power curve, but in this respect normal aircraft engines do not differ materally from one another.)

In a preferred arrangement according to this invention wherein the engine is coupled to a fixed pitch airscrew, the gear ratio of the rotor starter is such that when the engine is running at about half its normal ground revolutions the rotational speed of the rotor is between 40% and 60% in excess of its mean normal autorotative speed in flight at ground level, the power absorbed by the rotor at such an excess speed of rotation with the rotor blade pitch angle reduced substantially to zero being'about equal to the power available at half the normal ground revolutions of the engine.

It will be seen that with this arrangement a large excess of kinetic energy is stored in the rotor which is available for securing what may be termed a jump take-01f.

A still greater excess of kinetic energy may be imparted to the rotor, thereby giving a more pronounced Jump take-oil :by employing a controllable pitch airscrew for propulsive purposes.

If a controllable pitch airscrew is employed the gear ratio of the rotor starting transmission between the engine and rotor is preferably such that with the engine running at its normal (rated) revolutions the rotor is rotated at a speed between 50% and in excess of its mean autorotative revolutions in flight. When a controllable pitch airscrew is employed it is preferably ntrollable in accordance with the driving torque transmitted to the rotor, so that the airscrew pitch is decreased to a small value, preferably to zero, when a substantial torque is being transmitted to the rotor and increased to a value suitable for flight when the rotor driving torque vanishes.

For this purpose, a control connection may be provided between the rotor starter clutch and the airscrew pitch controlling means, whereby the airscrew pitch is decreased to a small value, preferably zero, on engagement of the clutch and increased to a value suitable for flight on disengagement of the clutch.

The arrangement is preferably such that when the clutch is disengaged the airscrew pitch may be independently controlled within a limited range of positive pitch variation to obtain the optimum airscrew efficiency in all conditions of flight.

As a further means of increasing the kinetic energy which can be stored in the rotor, suitable additional non structural masses may be incorporated in the rotor blades, preferably near their tips in order to increase the moment of inertia of the rotor as a whole. These additional masses are preferably located near the leading edge to bring the mass center f the blade forward for the purposes already stated.

Another advantage which accrues from the herein-described devices for controlling the blade pitch in accordance with the applied torque results from the fact that the application of a braking torque to the rotor automatically brings about an increase in pitch angle. This fact may be made use of for various purposes, for example to check the rate of descent when nearing the ground in a so-called vertical landing by partially applying the rotor brake whereby the pitch angle of the blades may be momentarily increased to check the descent just before touching the ground.

Other features of the present invention have reference to the undercarriage arrangements of the aircraft. In the first place the fact that during the application of the starting torque the rotor blades are substantially at no-lift pitch eliminates any overturning couple which might result from the lift of the rotor when the craft is on the ground, and if the air-screw is also arranged to deliver substantially no thrust during the running of the starter the overturning moment attributable to the airscrew thrust may also be eliminated. For this reason the main undercarriage wheels may be placed more nearly beneath the center of gravity than has hitherto been necessary, especially in aircraft of this kind in which the usual movable elevator is eliminated, the rotor axis itself being controllably tiltable for flight control purposes. By placing the wheels only very slightly in advance of the center of gravity an advantage is obtained in that the tail skid or tail wheel is less heavily loaded than with the wheels in an exaggerated forward position.

the full weight of the aircraft'is available for adhesion purposes and if further by controlling the airscrew pitch as above described the airscrew thrust is substantially eliminated during the running of the starter there will be substantially no tendency for the aircraft to run forward during the starting of the rotor and hence the wheel brakes may, if desired, be entirely eliminated. The wheel brakes are, however, preferably retained for parking or taxying purposes, but may be of much less capacity than hitherto usual, consideration being paid to the fact that in aircraft of the type herein referred to, being capable of landing with substantially no run, the use of wheel brakes for landing purposes is not essential.

The improved take-off" obtainable by means of the present invention also renders possible certain improvements with respect to aircraft adapted to fly on and ofi water. In this case the vertical or practically vertical take-off renders it unnecessary to provide the hull, float or floats with hydroplaning means, e. g. planing bottoms and steps, and in the case of amphibian aircraft the wheel undercarriage need not be made retractable as the water drag of the wheels becomes In the second place, the pronounced backward inclination of the rotor axis which is required for take-off" purposes with a fixed pitch rotoris no longer necessary in an aircraft having means according'to this invention enabling a substantially Vertical standing take-off to be made, so that the undercarriage (or flotation gear), body and rotor supporting structure may be so constructed and arranged that when the aircraft is standing on the ground (or riding on the water) the inclination of the rotor axis with respect to the vertical is of the same ordr as in normal horizontal flight. Thus also, wind gusts, when the machine is resting on the ground or on the water, have little or no effect.

Since with the arrangements of the present invention there is no tendency for the rotor to lift the aircraft before the starter is disconnected,

quiteunimportant. In general, the flotation gear may now, for the first time, be made substantially truly aerodynamically streamlined.

In aircraft of the kind referred to it has hitherto been usual toset the rotor axis so that an (imaginary) upward prolongation thereof is laterally inclined towards the retreating rotor blade, to compensate the angular deviation of the rotor thrust line from the rotor axis in the contrary direction in normal flight with horizontal translation. If the rotor axis is controllably tiltable it is usually mounted so that the above mentioned lateral tilt is present when the controls are held central laterally. The result of this is that when the rotor exercises lift without horizontal translation as in the socalled "vertical take-ofi," its thrust, which acts along the rotor axis in this case, applies an overturning couple to the aircraft in a sense tending to depress it on the side of the retreating rotor blade. For this reason the direction of rotation of the airscrew and rotor are preferably so chosen that they are the same when viewed from behind and above respectively so that the airscrew reaction torque tends to compensate the above-mentioned overturning couple of the rotor.

Other mechanism might be utilized to accomplish the foregoing objects, and it should be noted that the invention broadly contemplates a novel method of operating an aircraft having rotary sustaining blades, comprising initiating rotor operation at greater than normal R. P. M. and with the blades at less than normal pitch, effecting take-oil without power drive and with the blades at an increased pitch, and maintaining flight with the blades operating at their usual autorotative R. P. M. and pitch.

How the foregoing objects and advantages, together with others which are incident to the invention or will occur to those skilled in the art, are attained, will be clear from the description as a'whole, taken together with the accompanying drawings, in which drawings- Flgure 1 is a side elevational view of an aircraft embodying the present invention, with certain parts broken away;

Figure 1a is a diagrammatic illustration of the operation of one of the primary features of said aircraft;

Figure 6 is a view, in plan, of the rotor blade mounting mechanism of Figure 7, with the rotor hub shown in section substantially along the line 6-6 of Figure 7', and omitting the damping device and other parts adjacent the top of the rotor hub;

Figure 7 is a view, partly in section and partly in side elevation (further illustrating the structure of Figures 4 to 6). showing the upper portion of the rotor hub of the aircraft, and illustrate ing theattachment of the root end of one of the blades and the means for controlling certain blade movements;

Figure 8 is a detail view, partly in section, taken substantially along the line 88 of Figure '7;

Figure 9 is a somewhat diagrammatic side elevational view, illustrating in greater detail some of the control mechanism seen in Figure 1;

Figure 10 is a plan view of one of the rotor blades, partly sectioned near the tip;

Figure 11 is an illustration, partly in outline, somewhat similar to Figure '7, of a modification of the invention, and showing partly in elevation and partly in section a mechanism for locking the rotor blades as to certain pivotal movements thereof;

Figure 12 is a detail view of a portion of the modification of Figure 11, being taken substantially along the line l2l2 of that figure;

. Figure 13 is a somewhat diagrammatic view in side elevation, showing the mechanism for releasing the locking device embodied in the modification of Figures 11 and 12, and illustrating the operation thereof by the main control column of the aircraft;

Figure 14 is a diagrammatic view in side elevation of an alternative independent release for said locking device;

Figure 15 is a side elevational view, similar to Figure 1, but illustrating a modified form of the aircraft, particularly as to the undercarriage; and

Figure 16- is a fragmentary front elevational view of the lower part of the aircraft of Figure 15.

By reference now to Figures 1 to 3 inclusive, it will be seen that the aircraft, in general, comprises a body or fuselage l1; an engine l8 driving a tractor propeller or airscrew l9; and tail surfaces, including vertical fixed fins 20, 2|, and a horizontal stabilizer 22 having upturned tips 23 which contribute both to lateral and directional stability and have a cooperative action in conjunction with the control of the craft, as will further appear hereinafter. The aircraft also includes an undercarriage which may comprise a pair of telescopic shock absorbing struts 24, braces or radius rods 25, main wheels 26, and a steerable tail wheel 21.

The major means of sustension and control for the aircraft is the rotor, which comprises in general an upright hub or axis member 28 and a plurality (preferably three) of autorotative wings or blades 29, the rotor being mounted above the body by means of a pylon whichmay consist of struts 30, supporting at their apex the fixed spindle, axis, or bearing member 3| about which the rotor hub 28 rotates.

The power drive transmission for imparting the starting torque to the rotor comprises a horizontal shaft 32, preferably driven by the main propulsivev engine l3, and an upright shaft 33 which may be slightly inclined with respect to the vertical. One or more flexible or universal joint devices, or other means of flexibility, are preferably inserted in each of the shafts, as indicated at 34, 35.

'The rear end of the first mentioned shaft 32 and the lower end of the upright shaft 33 may be supported in any suitable bearings (not shown) in the casing 36, which contains a pair of bevel gears for interconnecting said shafts and a controllable clutch device for engaging and disengaging the drive connection, the clutch controlling lever being indicated at 36a. It is not necessary to illustrate the details in the said gear box or casing 35, since the gears and manually operable clutch may take the form of corresponding parts found in my prior Patent No. 1,673,232,

or may be patterned after rotor starter gears and clutches as now known commercially in this particular art, or as illustrated, for example, in the copending application of Joseph S. Pecker,

Serial No. 512,383, filed January 30, 1931, issued v April 30, 1935, as Patent #l,999,636.

The final connection, from the shaft 33 to the rotor hub 23, may be by another pair of bevel gears 31', 38, mounted respectively on the upper end of shaft 33 and on the rotor hubmember 28. The ratio of these gears should be so chosen that they, in conjunction with the bevel gears in the gear box 36, give the desired ratio between the prime mover and the rotor, as hereinbefore discussed.

The rotor head assembly preferably also includes a brake for stopping the rotor, The details thereof need not here be shown, as such brake may be similar to that illustrated in my issued Patent No. 1,948,456, or may take the commercial form now known in this art, or some form such as shown in the copending application of Joseph S. Pecker, No. 545,451, filed June 19, 1931, issued April 30, 1935, as Patent #l,999,737. The rotor brake operating lever, however, is shown at 39, and it should be understood that the brake, in this instance, has certain 'novel advantages, when used in conjunction with the automatic pitch' changing mechanism of the present invention, for example, to increase the tendency for the blades to go to a high pitch whenever that is desired, for. instance, to reduce the vertical speed of descent at the moment of landing, as hereinbefore described.

The control of the aircraft in fiight is obtained by tilting the rotor axis or bearing member 3| about a transverse pivot 40 and about a longitudinal pivot 4|; the member 3| being provided with a longitudinal lever 42 and a lateral lever 43 for control purposes. Longitudinal control of the craft is obtained by rocking the pilot's control lever 44 in a fore and aft direction, the motion being transmitted by a link 45, bell crank 46, approximately vertical push and pull rod 41, and the longitudinal control lever 42. Lateral control is effected by rocking the control stick 44 sidewise, the motion being transmitted by a rocking shaft 48, crank 49, substantially vertical rod 50, and the lateral control lever 43.

In Figures 1 and 2, the blades 29 are shown at rest, in thefull-line positions. With the rotor at rest, the longitudinal control movements are indicatedin dot-and-dash lines (Fig. 1) at 29a for forward tilt and at 29b for rearward tilt, of the rotor axis. The tilting for lateral control is indicated (Fig. 2) at 29c and 29d. These movements, in flight, control the craft (in conjunction with the weathercock action of the body and tail surfaces hereinbefore described) in the manner more fully set forth and claimed in my copending application No. 645,985, filed December 6, 1932, by shifting the lift line of the rotor (which is also disclosed in my British Patent 393,976). Another way of shifting said lift line is by pitchvarying mechanism of the character disclosed and claimed in my copending application No. 698,372, filed November 16, 1933 (disclosed also in my British Patent 410,532). 7

It should be clearly understood, however, that in flight the blades assume a coned position relative to the hub (as diagrammatically indicated by dotted lines at 29 in Figs. 1 and 2), this coning taking place individually as to each blade on its horizontal or flapping pivot, the axis of which is indicated by the reference character y. The dihedral or coning angle of each blade varies, as is now well known in this art, in accordance with variations in lift and centrifugal force and to eliminate detrimental gyroscopic effects. The coned position of the blades in flight also influences the magnitude of the pitching couple arising from the relative locations with the inner end of the member 59 hereinafter described.

Theintermediate member of the rotor blade articulation, hereinafter referred to as the drag ends of the plates 60, 6| are provided with integrally forged bosses 64, 65, the boss 64 being counterbored at 65 to receive the projection formed on the boss 65. The bosses are centrally traversed by a securing bolt 61, carrying a nut 68, and together form a very sturdy horizontal blade articulation pivot pin, rotatable inthe bore 56 of the lug or ear 55; the pivot axis y-y being the axis fOr blade coning and flapping.

The outer end of the forging 59 is formed as a stub axle, or secondary articulation axis, 69, on which are mounted the inner races of two conical roller bearings 10, the outer races of which are carried in a barrel-shaped housing II, the outer end of which is formed with a flange 12, to which is secured by bolts 13 the flange I4 of a socket member 15, forming the root member or attachment fitting'of the rotor blade. The rotor blade spar tube, which is fitted into said socket, may be of the center of mass and center of pressure of f amgle of coning of the blades on the hub; and it should be understood that the various dot-anddash line showings in Figures 1 and 2 are therefore only diagrammatic and for illustrative purposes.

In the embodiment of the invention here under discussion the airscrew or propeller is preferably of the controllable pitch type. Any suitable pitch controlling mechanism may be incorporated in the propeller hub 5|, the details of the device for varying the pitch of the propeller blades being not illustrated, as they may take any of the known forms, of which there are now several, such as the Hamilton-Standard, the Smith, and the Universal. The pitch controlling mechanism may be controlled by the collar shown'as associated. with the pivoted lever 52, to which lever is connected a rod 53 leading to the pilot's cockpit 54. The. pilots control connections for varying the airscrew pitch and for operating the rotor starter clutch device are illustrated in a purely diagrammatic form in Figure 1, and will? be hereinafter described more fully with reference particularly to Figure 9.

The rotor head assembly will now be described, as illustrated in detail in Figures 4 to 8 inclusive. The rotor hub member 28 preferably consists of a forging, the upper part of which is formed with three lugs or ears 55, which are bored at 56 to form or to receive the bearings for the horizontal or flapping articulation pivots. respectively, of the three blades of the rotor. The ears 55 also incorporate projections 51 and 58 which act as stops for limiting the upward and downward movements, respectively, of the blades on their horizontal articulations, by engagement secured therein by means of suitable studs or rivets 16. In order that rotor blades of different sectional contour and other characteristics may be applied to a given hub, for example for use on different machines, the socket 15 may be arranged to be variously set at different fixed angular positions about its own longitudinal axis; and for this purpose I may provide somewhat elongated apertures l3a where the bolts 13 pass through the flanges 12 and 14. In other words, the bolt holes may be somewhat oval in cross section, or elongated in a direction extending circumferentially of said flanges. In such event, it

is essential that the bolts be secured very tightly,

rected flange I1, to which is secured, as by studs or bolts 18 a retaining ring 19, the inner margin of which is flanged to engage with the outer race of the inner of the two roller bearing sets 10. The flanged ring 19 thus serves to transmit the centrifugal force of the blade to the inner bearing 10. From thence the forces are transmitted through the members 10a, 10b, 10c and nut 10 to the drag pivot member 69.

The axis of the drag pivot member or stub axle 69 is indicated by the line AA and the longitudinal axis of the rotor blade by the line B--B. It will be seen that the axis A-A is inclined upwardly and outwardly with respect to the axis BB, at an acute angle; said angle, in the example illustrated, being about 26". This angle, of course, remains the same regardless of the degree of coning or flapping of the assembly about the primary articulative axis yy. The pivotal connection 69, 10, H constitutes the drag pivot of the rotor blade articulative mounting, and it will be seen that when the rotor blade moves on its drag pivot, the blade axis B-B describes a conical path about the drag pivot axis AA, with the result that movement on the drag pivot is associated with a change of rotor blade pitch angle in such a way that when .the rotor blade moves on'its dragpivot in a general direction contrary to the direction of rotation of the rotor the pitch angle is decreased, whereas when the rotor blade moves on its-drag pivot in a direction generally forwardly with respect to the direction of rotation ofthe rotor, the pitch angle of the blade is increased. The direction of rotation of the rotor is indicated by arrows in Figures 3 and 6.

Movement of the blade about its drag pivot is limited by a stop plate 88 secured as by bolts -8I to the retaining ring I9 andhaving a pair of inwardly projecting horns 82, whose inner faces 82a and 82b are adapted to engage with the flat i faces 59:: of the inner part of the stub axle forging 59. The stops 82 are'so positioned that is approximately zero, whereas with 'the rotor blade in its most forward position the blade pitch' angle is approximately 8", or slightly less, the

blade pitch angles being taken as measured from the angle of zero lift with respect to a plane perpendicular to the axis of rotation of the rotor .as a whole. It should be understood that there is no necessity for a range of pitch extending above t'ie autorotational range and into the high pitch range heretofore required in helicopter design, since the rotor of the present invention, when the blades are thrown into their pos- 'itive pitch, is acting upon undisturbed air, in

illustration in Figure la.

contrast to the high slip-stream velocity in which helicopter propellers must operate.

The conical movement of the blade, and the preferred range of pitch angl variation will be very .clearly understood from the diagrammatic In that figure, 29c represents the radial or mid-position of a given section of a rotor blade, taken at some distance out from the root end. the dot indicating the center of the blade spar. At 29} the blade is moved to its extreme position of lag, where the blade pitch angle is zero: and at 29g the blade is moved to its extreme position of lead, where the blade pitch angle is 8. In this figure. the arrow indicates the direction of rotation of the rotor and also represents a plane perpendicular to the axis of rotation, and it will be observed that when the blade is in mid-position, it is approximately at 4 with respect to said plane. If the blade is designed to assume a position of equilibrium which is somewhat lagging, it will be understood that the normal pitch angle, in flight, will be somewhat less than 4". It should also be understood that this figure does not indicate the degree iof'coning angle which the blade will take about its horizontalpivot, when under flight load. Furthermore, a symmetrical section is shown in the figure, although it will be understood that for a ran e of ,pitch angles from to 8, a higher lift/ drag blade section would normally be employed,

the showing in this figure being purely diagram matic. I

When the starting torque is applied by means of the power transmission, the inertia of the rotor blades causes them to .swing back on their drag pivots until the stops 82a of the stop plates 88 engage the forward flat faces 59.1: of the member 59, as shown in dotted lines in Figure 8. In this 'position the plate (and thus the blade) is restrained from rotation on the drag pivot by a friction device comprising a plunger 83 slidable in the pocket 84 mounted on the retaining ring I9. The plunger is pressed inwardly by a powerful spring to engage frictionally with the rounded face 592 of the member 59, the spring 85 being retained by an adjustableabutment screw 88. 'The plunger may be faced with frictional material at 83c, and the frictional face as well as the face 592; may be so relatively curved or inclined that, as the face 59:: approaches the stop face 82a, the spring 85 is more strongly compressed, whereby the frictional restraint is greatest when the blade is moved to its fullest extent in the backward or lagging direction, as limited 'by the stop face 82a. Further, the friction plungas might happen if the engine choked or misfired. Such momentary reduction of torque might cause the blades to swing forward on their drag pivots andincrease their pitch angle, and might therefore result in the aircraft leaving the ground before it was intended to do so. The above-described arrangement is therefore designed to ensure that the rotor blade pitch angle is not increased from the minimum value until the starting torque is definitely destroyed, as by disengaging the starter clutch.

The movement of the rotor blades on their drag pivots is damped by means of a central friction damper mounted on the hub member 28, this-device being illustrated in Figures 4, 5 and 7. Each blade is connected with the friction damper by an independent linkage, consisting of a forked hinge plate 88 secured on the three uppermost of the studs I8 and carrying a hinge bolt 89, on which is pivoted a link 98 terminating in a ball 9| whichis engaged on either side by a pair of set screws 92 secured in a slotted and forked link 93 and locked by locking nuts 94. The inner forked end of the link 93 is pivotally connected with a hinge plate 95 by means of the bolt 98, the hinge plate being secured by rivets 91 to a projecting part 98 of a flat ring 99 forming one of the movable elements of the central friction damper.

The latter device comprises a support I08 screwed into the top of the hub forging 28 and carrying a bell-shaped member IN, the lower part of which is outwardlyl flanged to form a friction pressure washer I82, and on which are independently rotatable three of the said flat rings 99, which are spaced from one another by friction washers and each connected with one of the rotor blades by means of a linkage as just described.

The upper end of the bell-shaped member I8I terminates in a threaded sleeve I83 carrying nuts I88 by which pressure can be applied through a spacing washer I85 to an upper pressure washer or cap I88, which transmits pressure through a cushion ring or washer I81 to the friction rings 99 which are thus held between the upper and lower pressure elements I86 and I82, the friction between them being variable by interconnect the central damping device with the a three drag pivot (secondary articulation) housings, permit free movement of the rotor blades on their flapping (primary) pivots, but serve to transmit movements of the blades about their drag pivot axes A--A to the friction damping rings, the damping being effective with regard both to symmetrical blade movements relative to the rotor hub and to unsymmetrical movements of the blades relatively to one another; the dot and dash lines showing, in Figure 7, illustrating the freedom of movement of the link by which is accommodated the coning and flapping of the blades on their horizontal pivots.

Turning now to the control devices, it will be seen that Figure 9 shows the connections for controlling the clutch of the rotor starting mechanism and the pitch of the controllable pitch airscrew. The rotor starter clutch controlling lever 36a is biased by a spring I08 to a clutch-disengagin p ition. A rod I09 connects the lever lid with acontrolling lever IIII in the pilot's cockpit.

The airscrew pitch controlling lever 52 (se Figure 1) is connected by rods III, II2 with an airscrew pitch controlling lever H3 in the cockpit. Rod III is provided with a slot Ill engaging a pin II! mounted on the clutch control lever III), and is pivotally connected at IIB with said rod I I2, to allow for varying angularity with respect to the latter rod. It will be seen that movement of the clutch control lever III! will be transmitted to the airscrew control rod III, but the pin and slot connection II, II5, allows a limited movement of the airscrew control II3 independently of the operation of the clutch control. In the full line position of the various levers and associated parts, in Figure 9, the rotor starter clutch is disengaged and the airscrew pitch control is at that point which sets said pitch at the lower limit of the range of variation useful in flight. With the clutch lever in this position, the airscrew control lever can be thrown forward to the position indicated at I I3a, which brings the airscrew pitch up to the ma'xi-. mum value useful in flight. Upon engaging the clutch by throwing the lever IIII into the position shown in dotted lines at IIIIa, the airscrew control lever H3 is brought back into the position indicated at II3b, in which position the airscrew pitch is decreased to a very small value, preferably to zero.

Turning now to the construction of the rotor blade itself, it will be seen in Figure that the blade 29 is constructed with its mass center M closer to the leading edge I" then is the line 21-11. The line H (shown in dot and dash lines in Figure 10) lies at a distance of about 25% of the chord dimension from the leading edge I" of the blade; and represents that line with reference to which the pitching moment coeficient of the blade is substantially constant. To bring the mass center of the blade sumciently far forward, and to increase the total moment of inertia of the rotor in an economical manner, small weights'm are removably mounted in the forward outer portion of the blade, as by means of studs II8 which are screwed into the collar devices H9 secured to the main rotor blade spar or tube I20. Adjustment of the weight as well as of the sectional and longitudinal location of the center of gravity of the blade may thus also be obtained.

The direction of rotation of the rotor is indi cated in Figure 10 by an arrow, as well as in Figures 3 and 6. The direction of rotation of the propulsive screw is indicated in Figure 2. By comparison of these figures it will be seen that the direction of rotation of the airscrew, as viewed from behind, is the same as the direction of rotation of the rotor as viewed from above; the reason for this arrangement having ing couple imposed by the offsetof the rotor lift line from the longitudinal vertical plane through the center of gravity of the craft as a whole, at the moment of vertical take-off, practically neutralize each other.

Turning now to Figures 11 and 12, it will be seen that I have shown a modified arrangement for retaining the rotor blades in their minimum pitch position while revving up the rotor for take-off. This arrangement incorporates a positive lock for the blades, which comes into action automatically when they reach the limit of their movement, about the drag pivots, in the backward direction; that is, when they are in the minimum pitch position. In this arrangement, there is a little cylinder or pocket 84a (similar to that shown in Figure 8), containing a plunger 831), which is urged by a very light spring a; and in this case the plunger has its extremity formed as a tooth 830 which engages in a notch I2I formed in the rounded face of the drag pivot member 59. When the stop face 82a comes up against the surface 59m, the tooth 830 is pressed into the notch I2I, as indicated in dotted lines in Figure 12.

The engagement of thetooth with the notch takes place automatically under the action of the spring 85a, when the rotor blade has reached the limit of its backward movement, and the blade is then positively locked until the tooth is released from said notch. This is accomplished by means under the control of the pilot, as follows:

The plunger 83b is provided with a projection I22 which passes through a slot formed in the side wall of the cylinder 84a and which is engageable, when the blade is at the limit of its backward travel about the drag pivot, by a button I23 (see Figure 11) formed on the upper end of a rod I24 which is slidably mounted in the forging 59. The lower end of the rod I24 carries a roller I25 which is engageable by a horizontal ring I26 supported on three vertically slidable rods I21 mounted in brackets I28 secured to the rotor support member 3|. Pivoted on the brackets I28 are ,levers I29, the inner ends of which engage the lower ends of the rods I21, and the outer ends of which are connected to Bowden type control cables I30, the tension of which is maintained by cushioning springs I3I which seat on projections I32 of the brackets I28; the projections forming the abutments of the sheaths I33 of the Bowden cables I30. The three cable devices I30 are brought together at a point below the rotor and coupled to a single Bowden control I34 which is led into the pilot's operating the control cable I34. A preferred arrangement is to connect this cable to a downward extension of the rotor starter clutch control lever IIIJ, as seen in Figure 9. This operates ring I 23, the motion being transmitted by. the

rollers I to the rods I 24 and projections I23 so as to lift the plungers I22 against the pressure of the springs 85a and disengage the teeth 830 from the notches I2I, thus allowing the rotor blades to swing forward on their drag pivots and increase their pitch angle.

An alternative control is illustrated in Figure 13, in which the Bowden cable I34 is attached to a projection I35 formed on a plunger I36 which is mounted in a fixed pocket I31 and loaded by a spring I38, which latter forces the plunger I36 in a direction for tensioning the cables I34, I32, thereby releasing the blade locking device. The plunger I36 is engageable by the main control stick 44, in the cockpit. When the latter is anywhere within the range of positions utilized in normal flight, there is a clearance between the control column and the plunger I36, the limits of the flying range of control column positions being shown by the dotted line positions 44a and 44b, in Figure 13. When the control column is moved forward well beyond the flying range it engages the plunger I36 and moves it forward against the spring I31 so that the Bowden" controls allow the levers I29 to rock and the members I21, I24 and I22 to drop downwardly, with the result that the plungers 8312 are no longer restrained against their spri gs 85a; and upon a swinging of the rotor blades into their rearmcst positions (when the starting torque is applied) they are thereupon locked by engagement of the teeth 830 with their notches I 2I. On bringing back the control column from its extreme forward position, a release of the locking devices takes place.

In Figure 14, I have illustrated a modification in which the Bowden control cable I34 is operated by an independent control consisting of a small hand lever I 39 pivoted on a bracket I40 which may be mounted, for example, on the instrument board MI in the pilots cockpit 54. Release of the locking device takes place On moving the lever I39 downwardly. It will be noted that unless the locking device is released it cannot engage again in normal flight, as the rotor blades do not then swing backwards on their blade pivots far enough to permit the teeth 830 to engage the notches I2 I.

Turning now to Figures 15 and 16, it will be seen that I have illustrated a modification of the aircraft previously described; this modification incorporating an amphibian landing gear. This arrangement incorporates twin floats I42 supported on a structure of struts I43 and interbraced by transverse members I44; and within each float is housed a main supporting wheel I45, the wheels being located within recesses in the floats. These wheels may be of the air cushion type, as indicated, or may be yieldably supported on any suitable shock absorbing devices. The wheels may permanently project slightly below the bottoms of the floats; the floats, in general, being constructed of a substantially streamline form, from an aerodynamic standpoint, and without the usual hydroplaning means, such as planing bottoms or steps. The construction of the floats in this manner, and the permanent location of the landing wheels so that when the craft is on-the water the wheels are practically submerged, is rendered possible by the capability of the machine to make a substantially vertical take-oil; and it is therefore possible to utilize floats with a substantially streamline form, as above mentioned.

It may also be noted in connection with this machine that I contemplate a fixed pitch propeller I3a, and a choice of gearing in the rotor starter, including the gearing 31a, such as to a make possible a preliminary turning up of the rotor to about 40% to 60% above the normal autorotational speed thereof, when the engine I8a is running at about half its normal ground revolutions; for the reasons hereinbei'ore set forth.

With further reference to the undercarriage and float arrangements, not only as illustrated in Figures 15 and 16, but also as illustrated in Figures 1 to 3, it will be seen that the main wheels are placed only. slightly in front of the center of gravity of the aircraft when the latter is at rest; the center of gravity being indicated at g in Figures 1, 2 and 15. This location of the wheels is made possible by virtue of the character of the take-off and landing accomplished by the present invention, and is desirable in reducing the loads upon the tail wheel which would otherwise be quite heavy in view of the manner of controlling this type of craft by means of tilting the rotor. I am thereby also enabled to make the fuselage construction lighter, toward the rear end, than would otherwise be the case.

Furthermore, theundercarriage is made as low as possible, consistent with adequate clearance for the propeller. It should also be noted that the disposition of the aircraft body and rotor supporting structure with respect to the wheels and/or floats is such that when the aircraft is at rest on the ground or on the water the rotational axis of the rotor, indicated by the dot and dash line :c--.'r, is only slightly inclined backwards from the vertical.

Referring back to the matter of the automatic change of rotor blade pitch angle, it will now be clear that said pitch angle, being dependent on the position with each blade takes with respect to its inclined drag pivot axis A-A, is variable in accordance with the centrifugal force experienced by the blades as well as the torque transmitted to the rotor hub. The relation between change of rotor blade pitch angle and angle of lag (as hereinbefore defined) depends on the relative angular setting of the axis A--A of the drag pivot pin 69 and the axis 3-3 of the blade. The angular setting herein shown, namely 26, has been found suitable to give the required relation between variation of centrifugal force experienced by the blades and variation of blade pitch angle with blades whose rotational speed and moment of inertia are of the order now in common use. The angular setting of the drag pivot pin may, however, be varied within fairly wide limits, to conform with the design requirements of the particular aircraft and particular rotors.

In considering the relativeR. P. M. of the engine and the rotor, upon starting of the latter prior to vertical take-off, it might be mentioned that various rotors at present in use are designed to autorotate in normal flight at say R. P. M., R. P. M., or 1'75 R: P. and some may be designed to operate even faster than that. It

being different, in the manner hereinbefore set forth, in a machine having a controllable pitch propeller as compared with one having a flxed pitch propeller. By way of example, with a rotor designed to normally autorotate at 150 R. P. M. and an engine designed to have a rated R. P. M. of 2000, if a controllable pitch propeller is employed, the ratio of the rotor on should be such as to give a rotor R. P. M. somewhere between 225 and 300, when the engine is operating at its rated R. P.-M.

In conclusion, an aircraft in accordance with the present invention is not only capable of the usual translational flight and vertical descent characteristic of this type of machine, but also embodies the capability of a substantially vertical take-01f (on an approximately even keel) while eliminating the two major drawbacks characteristic of various helicopter types of machine which have been heretofore experimented with, namely: the reactive torque of the lifting means during the vertical take-oil, and the difllculty of controlling the machine during vertical takeo-if or such steep take-off as would render the normal type of control surfaces ineifective. These results are obtained by virtue of the combining of the stored kinetic energy system for take-oil and the rotor tilting system by which the control of the machine is placed directly in the rotor, the rotor being effective for control purposes, even without any appreciable movement of the body of the machine. Furthermore. this combination of arrangements, in conjimction with the propulsion means, and especially with a variable pitch propeller, makes possible a quick transition from the vertical take-oil into the forward translational movement without appreciably losing the altitude gained in the takeoff.

The terms rotor, air rotor, "sustaining it should be pointed out that of rotational speed to generate temporarily a thrustsufllcient to raise the aircraft substantially vertically from the ground, said last mentioned means including mechanism automatically operative in response to variations of the torque applied to the rotor to decrease the mean pitch angle of the rotor blades during the application of the starting torque to a value substantially less than that normally employed in flight and to increase said mean pitch angle on the vanishing of said starting torque at least to the minimum value employed in normal flight .operation.

3. An aircraft as claimed in claim 2, characterized in that the application of a driving torque to the rotor, at least with the rotor turning at less than the minimum autorotational speed of flight, effects a mean pitch.

angle setting of the rotor blades of substantially zero value, i. e., the value for minimum aerodynamic rotational drag. I

4. An aircraftas set forth in claim 2, and in which the automatic rotor blade pitch changing means are responsive also to changes of centrifugal force experienced by the blades, ran increase of centrifugal force tending to increase the mean pitch angle and a decrease tending to decrease it; the construction further being such that the influence of the torque transmitted to the rotor so preponderates over the influence of centrifugal force that the application of maximum torque suffices to maintain the mean blade pitch angle at or near its minimum value-at least while the rotational speed of the rotor does not exceed the minimum autorotatlonal speed of fli ht.

5. For an aircraft, a sustaining rotor comprising a generally upright hub structure, a plurality of blades or wings located thereabout in position to be normally autorotated by the relative flight wind, means providing for independent swinging of each blade relative to the hub member in a general up and down or flapping rotor, lifting rotor, and the like, unless otherwiselimited by the context, are to be understood as inclusive of thrust-producing air rotors in general, whether or not power driven.

I claim:

1. In an aircraft, an autorotatable sustaining rotor comprising a flappingly pivoted aeroform blade having its mass center located forwardly of its sectional center of pressure and spaced outwardly at least half the radius from the center of rotation, means for driving the rotor, and means for automatically varying the effective pitch of the blade under the influence of the driving means.

2. In an aircraft having means of propulsion and having a sustaining rotor comprising wings or blades mounted in position to be normally autorotationally actuated in flight, a power transmission mechanism for app g a driving torque to the rotor to impart thereto an initial rotation prior to flight and including means providing a transmission ratio such as to impart to the rotor an initial rotational speed considerably in excess of the mean autorotative speed of the rotor in flight, and means for sub-- stantially decreasing the rotational drag of the rotor during the application of the starting torque and for utilizing the excess kinetic energy stored in the rotor by virtue of its initial excess direction and including a pivot having its axis substantially in a plane perpendicularly to the rotational axis of the hub, means providing for independent swinging of each blade relative to the hub member in a general direction fore and aftin the rotative path and including a pivot having its axis substantially in a plane contain- Y ing the rotational axis of the hub, an automatic means for decreasing the pitch angle of each rotor blade responsively to its movement on the latter pivot in a direction contrary to the direction of rotation of the rotor.

6. A rotor as claimed in claim 5 and wherein the second mentioned pivot is inclined upwardly and outwardly with respect to the longitudinal axis of the blade, 1. e., so that the upper angle bounded by the longitudinal blade axis and the said pivot axis is less than a right angle.

'7. A rotor constructed in accordance with claim 5 and including means for damping movements of the blades about the second mentioned pivot.

8. In an aircraft, having a sustaining rotor including a plurality of pivoted rotor blades, means for applying a starting torque to said rotor, means for imposing a braking effect upon said rotor, and means for varying the pitch of the rotor blades automatically responsive to decrease their pitch under the influence of the starting torque and to increase their pitch under the influence of the braking force.

9. In an aircraft, an autorotatable sustaining rotor comprising a plurality of aeroform blades mounted to provide for lag anl lead movements with respect to each other, mechanism for applying a driving torque to the rotor, and mechanism operative under the influence of the driving torque for lowering the mean eflective pitch of all the blades of said roto 10. In an aircraft, an autorotatable sustaining rotor comprising a plurality of aeroform blades mounted to provide for lag and lead movements with respect to each other, mechanism for applying a driving torque to the rotor, and mechanism automatically operative by cessation of the driving torque to raise the mean effective pitch of all the blades of said rotor.

11. In an aircraft, an autorotatable sustaining rotor comprising a plurality of aeroform blades, a power plant with a disconnectible driving system coupled .to said areroform blades, and mechanism automatically operative by disconnection of said system to raise the mean effective pitch of all the blades of said rotor.

12. In an aircraft, an autorotatable sustaining rotor comprising a plurality of aeroform blades, pivot means for each blade providing for individual swinging movements of the several blades under the influence of flight forces thereon, a disconnectible power drive system for a plying a torque to the rotor, and an additional pivot device for each blade providing for blade movement between a low pitch position and a higher pitch position in a path such that the centrifugal force of rotation of the blade acts to move theblade from said low pitch position toward the higher pitch position upon disconnection of the drive.

13. In an aircraft, an autorotatable sustaining rotor comprising a plurality of aeroform blades, pivot means for each blade providing for individual swinging movements of the several blades under the influence of flight forces thereon, a disconnectible power drive system for applying a torque to the rotor, and an additional pivot device for each blade providing for blade movement between a low pitch position and a higher pitch position in a path such that the centrifugal force of rotation of the blade acts to move the blade from said low pitch position toward the higher pitch position upon disconnection of the drive, and means for positively retaining all the blades in a low pitch position during operation of the driving system.

14. An autorotatable rotor including a hub and a plurality of sustaining blade elements movably mounted to compensate for differential lift forces, and for each such blade element a pivot device for varying the effective pitch setting thereof and providing for change of effective pitch from a positive value to approximately zero value, the blade element being so positioned with relation to the pivot that whenever the blade element is near zero effective pitch setting and the rotor is turning by virtue of its momentum the action of centrifugal force on the blade is continuously urging it to move on its pivot to a position of positive efiective pitch, and releasable mechanism for positively holding all said blade elements simultaneously at approximately zero effective pitch position against the action of centrifugal force of rotation.

15. In an aircraft, a normally autorotationallyactuated sustaining rotor comprising blades articulated individually for oscillation in flight, disconnectible means for applying a driving torque to the rotor, and mechanism for varying the mean pitch of tlie rotor blades including means for lowering their pitch during application of the driving comprising an auto-rotational rotor for susten-' sion, said rotor including a plurality of sustaining blades, means providing for the transmission of torque from said power means to said rotor, mechanism providing for variation of the pitch angles of the rotor blades and reduction of said blade pitch angles to a value less than that normally utilised in autorotative flight during the application of the starting torque, and means ensuring that disconnection of the transmission means, when the rotor is rotating at a normal flight speed, is momentarily accompanied by an increase of bladepitch angle from said reduced value to a value greater than that normally utilised in autorotative flight.

17. In an aircraft having means of propulsion and having a sustaining rotor comprising wings or blades mounted in position to be normally autorotationally actuated in flight, a power transmission mechanism for applying adriving torque to the rotor to impart thereto an initial rotation prior to flight and including means providing a transmission ratio such as to impart to the rotor an initial rotational speed considerably in excess of the mean autorotative speed of the rotor in flight, and mechanism mounting the blades for movement including pitch varying movement for lowering the pitch of the blades and thus substantially decreasing the rotational drag of the rotor during the application of the starting torque, the blade mounting mechanism comprising means operative under the influence of centrifugal force on the blades to increase the pitch of the blades upon cessation of said torque whereby to utilize the excess kinetic energy stored in the rotor by virtue of its initial excess of rotational speed to generate temporarily a thrust suflicient to raise the aircraft substantially vertically from the ground.

18. In an aircraft having means of propulsion and having a sustaining rotor comprising wings or blades mounted in position to be normally autorotationally actuated in flight, a power transmission mechanism for applying a driving torque to the rotor to impart thereto an initial rotation prior to flight and including means providing a transmission ratio such as to impart to the rotor an initial rotational speed considerably in excess of the mean autorotative speed of the rotor in flight, and mechanism mounting the blades for movement includin pitch varying movement for lowering the pitch of the blades and thus substantially decreasing the rotational drag of the rotor during the application of the starting torque, the blade mounting mechanism comprising means operative under the influence of centrifugal force on the blades to increase the pitch of the blades upon cessation of said torque whereby to utilize the excess kinetic energy stored in the rotor by virtue of its initial excess of rotational speed to generate temporarily a thrust suilicient to raise the aircraft substantially vertically from the ground, said last mentioned mechanism providing for decreasing the mean pitch angle of the rotor blades during the application of the starting torque to a value substantially less than that normally employed in flight and for subsequently increasing said mean pitch automatic increase employed in flight operation.

opposite to the direction 12 angle at least to the minimum valuenormally 19. In an aircraft, means of sustension comprising a rotor having wings or blades mounted to provide for lag and lead movements relative to each other and arranged to be capable of autorotational actuation in flight, a power transmission mechanism for imparting a driving torque to the rotor, and means automatically responsive to changes both of torque and of centrifugal force experienced by the rotor blades to automatically reduce the mean pitch angle of all the rotor blades when the torque force predominates and to increase the same when the centrifugal force predominates.-

20. An ircraft including power means for forward translation, normally air actuated means comprising an autorotational rotor for sustension including at least one providing for the transmission of torque from said power means to said rotor, rotor blade mounting mechanism providing for variation of rotor blade pitch including means for lowering the pitch angle during the application of power thereto, the mounting mechanism further including means for subsequently raising said pitch angle automatically under the influence of centrifugal force on the rotor blade upon disconnecrotor blade, means Y tion of the power transmission whereby said rotor eifects vertical ascent, and means providing for free rotation of said rotor for vertical descent.

21. The method of operating an aircraft, having a forward propulsion power plant and a system of sustaining blades rotatably mounted for autorotative actuation by relative airflow, which comprises delivering power from said power plant to rotate said blades up to a speed substantially in excess of their normal autorotational speed, with said blades at a reduced pitch ineffective to raise the aircraft, maintaining said pitch substantially uniform as to the several blades during the application of power thereto, thereafter cutting oi the delivery of power to the blades and raising the pitch of the severalblades to an increased itch angle greater than the normal, so as to utilize the kinetic energy of the rotating blades in eifecting a vertical take-oflof the craft, and continuing the rotation of said blades, autorotationally, at a reduced speed and an intermediate normal pitch by eflecting a forward translational movement of the craft, and leaving said blades free for automatic individual pitch variation under the influence or the varying balance of aerodynamic forces occurring in forward fli ht.

22. In an aircraft, a sustaining rotor including a plurality of blades postioned to be normally autorotated about a common axis, an upright axis member for said blades mounted to be tilted in all directions for control purposes, a control member for tilting the rotor, said blades being pivotally connected with said axis member for oscillative movements relative thereto whereby the thrust line or the rotortends to shift laterally during forward translational movement of the craft, the axis of said axis member being set on said mounting at a laterally inclined angle toward which said thrust line moves during translational flight, the oil'set being such that with said control member held in neutral position the projection of said axis member is laterally toward one side of the center of gravity of the craft, power transmission means forimparting an initial rotation to said rotor prior to take-off, means for reducing the lift incidence, means said lift-line whereby to provide pitch angle of the rotor blades during the application of the starting torque and for increasing their pitch angle subsequent thereto whereby substantially vertical takeoff or the craft is obtained, and means of forward propulsion for the craft having a net reactive torque so arranged as to tend to counteract the rolling couple introduced by the lateral inclination of the rotor axis during vertical ascent.

23. In an aircraft, power means including an airscrew for forward propulsion, means of sustension comprising rotor blades adapted to be normally autorotationally actuated by the relative flight wind, disconnectible means for rotatively driving said blades by said power means prior to take-oil, mechanism pivotally mounting said rotor blades upon the craft for differential-lift compensation, for variation in incidence, and for shifting of position of the rotor lift-line relative to the center of gravity of the craft, means responsive to application of driving torque to said blades to position them at approximately zerofor raising the blade incidence to a substantial positive-lift position for direct or vertical take-oil, and controllable means operating upon asid blades to variably position control over the attitude of the craft and the angle of the direct take-oil when the drive means is disconnected and the blade incidence is raised.

24. In an aircraft, power means,means of sustension comprising rotor blades adapted to be autorotationally actuated by the relative flight wind at least when descending without power, disconnectible means for rotatively driving said blades by said power means, mechanism pivotally mounting said rotor blades upon the craft for differential-lift compensation, for variation in mean incidence and for shifting of position of the rotor lift-line relative to the center of gravity of the craft, means responsive to application of driving torque to said blades to position them at a reduced incidence ineffective to raise the aircraft, means for raising the blade incidence to a substantial positive-lift position for direct or vertical take-off, and controllable means operating upon said blades to variably position said liftline whereby to provide control over the attitude of the craft and the angle of the direct take-off.

- 25. In an aircraft having means of propulsion and having-a sustaining rotor comprising wings or blades mounted in position to be normally autorotatively actuated in flight, a power transmission mechanism for applying a driving torque to the rotor to impart thereto an initial rotation prior to flight and including means providing a transmission ratio such as to impart to the rotor an initial'rotational speed considerably in excess of the mean autorotative speed of the rotor in flight, mechanism for substantially decreasing the rotational drag of the rotor during the application of the starting torque and for utilizing the excess kinetic energy stored in the rotor by virtue of its initial excess of rotational speed to generate temporarily a thrust sufllcient to raise the aircraft substantially vertically from the ground, said mechanism including means automatically operative in response to variations of the torque applied to the rotor to decrease the mean pitch angle of the rotor blades during the application of the starting torque to a value substantially less than that normally employed in flight and to immediately increase said mean pitch angle, upon the vanishing of said starting torque, at least to the minimum value employed in normal movement of the blade with respect to flight operation, and restraining means operative during the application of the starting torque for preventing the pitch angle of the rotor blades from increasing to a value sufllcient to raise the aircraft from the ground as a result of an accidental momentary or transient diminution in the applied torque.

26. An aircraft in accordance with claim 25 and in which the restraining means are automatically engaged upon movement of the rotor blades toward their lower limit of pitch angle.

27. An aircraft in accordance with claim 25 and in which the restraining means include a frictional device operative automatically to resist change of rotor blade pitch angle at least when said pitch angle is near its minimum value.

28. An aircraft in accordance with claim 25 in which the restraining means comprise an auto matic locking device operative when the blades are approximately at their minimum pitch angle.

29. An aircraft in and in which the restraining means include a locking device and a manual control to release the same.

30. An aircraft in accordance, with claim 25 and in which the restraining means comprise a locking device for retaining the rotor blades at a low pitch position and an interconnection between said device and a control for the power transmission mechanism, whereby, upon disengagement of said transmission mechanism, the rotor blades are released for automatic movement to a position of increased pitch angle.

31. An aircraft in accordance with claim 25 wherein there is also a longitudinal flying control, and an interconnection between said control and said restraining means, whereby the restraining means is released when the fore and aft position of the longitudinal flying control is within the normal range of control movement employed in flight, but is rendered operative when the said control is moved forward well beyond the range of movement normally utilized in flight.

32. For an aircraft sustaining rotor, a rotor blade mounting pivot providing for swinging the rotor hub, means for effecting conditions of relative freedom for and restraint of said blade swinging movement, and mechanism for regulating said means as between said two conditions including an element manipulable by the pilot while the rotor is rotating and an actuating connection extended from said element to said means.

33. A construction in accordance with the preceding claim in which said means is constructed to effect said restraint by positive locking of the blade as against swinging movement.

34. For an aircraft, a sustaining rotor comprising an upright rotative hub or axis member, an elongated sustaining wing or blade positioned to rotate thereabout, means for pivotally securing said blade on said axis member including a flapping pivot with its axis located substantially in a plane perpendicular to the rotational axis of the rotor and an inclined drag pivot with its axis located substantially in a plane containing the rotor axis, the longitudinal blade axis and the axis of the drag pivot forming an angle less than 90 at the upper side of the blade and outwardly of the drag pivot, the mass distribution and aerodynamic characteristics of the blade being such that the blade in flight tends to take up a position of equilibrium in which the centrifugal accordance with claim 25 force exerts a couple about the drag pivot acting forwardly, i. e., in the direction of rotation, so that an increase of rotational speed causes the mean blade pitch angle to increase and vice versa.

35. Means for mounting a substantially radially extending rotor blade upon a substantially upright rotor hub, comprising a drag link pivoted to said hub on a horizontal axis, a secondary pivot pin mounted on said link with its axis inclined upwardly and outwardly from the axis of the hub, a frame or housing around said inclined pivot and connected with said blade, and antifriction' bearing means interposed between the inclined pivot and said housing.

36. Means for mounting a substantially radially extending rotor blade upon a substantially upright rotor hub, comprising a drag link pivoted to said hub on a horizontal axis, a secondary pivot pin mounted on said link with with its axis inclined upwardly and outwardly from the axis of the hub, a frame or housing around said inclined pivot and connected with said blade, and radial and thrust bearing means interposed between the inclined pivot and said housing.

37. A rotor blade articulation including a pivot providing freedom for blade swinging movements under the influence of acre-dynamic forces the axis of which pivot lies approximately in a plane containing the rotor axis and extends at an upwardly and outwardly inclined angle with respect to the major blade axis.

38. For an aircraft, an autorotatable sustaining rotor comprising a substantially upright rotative hub or axis member, an aeroform sustaimng wing or blade, means for pivotally securing said blade on said hub including a flapping pivot-axis located approximately in a plane perpendicular to the general rotational axis of the rotor and an inclined pitch-varying pivotaxis making an acute angle with the longitudinal axis of the blade and located approximately inclined pivot-axis cause the blade in autorotative flight to taken an average position of aerodynamic equilibrium in an intermediate region of the operating range of pitch angular move ment.

39. In an aircraft sustaining rotor, a rotatable hub, an elongated blade or wing, and mechanism mounting the blade on the hub including a plurality of pivots, one of said pivots providing for bodily swinging of the blade to accommodate variable flight loads, and another of said pivots providing for variation of the pitch of the blade and being. angled with respect to the longitudinal axis of the blade, and means automatically setting up a substantially greater but yielding resistance to pivotation of the blade on the latter pivot than on the former.

40. In an aircraft sustaining rotor, a hub, an autorotatable blade, and pivot mechanism securing said blade to said hub including a pivot owing to the obliquity of said pivot, and a device setting up resistance to pivotal movement of the blade about said pivot axis and located to act independently of said bearing means whereby the resistance of said device is unaffected by variations in thrust on said bearing means.

41. In an aircraft, a forward propulsion engine, a controllable pitch air screw driven thereby, a normally autorotative rotor, a disengageable transmission between said engine and the rotor for applying a starting torque to the latter, the ratio of transmission between the engine and the rotor being such that with the engine running atapproximately its normal (rated) R. P. M. the rotor is rotated at a speed substantially in excess of its 'inean autorotative R. P. M. in flight, and a clutch device for effecting said'engagement and disengagement of the transmission, and means whereby the air screw pitch is controlled in accordance with the driving torque transmitted through said clutch device to the rotor, so that the airscrew pitch isdecreased to a small value, approximately zero, when a substantial torque is being transmitted to the rotor and increased to a value suitable for flight when the rotor driving torque disappears.

42. An aircraft in accordance with claim 41, in which there is a manually controllable member for engaging and disengaging said clutch,

and a control connection between said member and the airscrew pitch controlling means, whereby the airscrew pitch is decreased to a small value, preferably zero, on engagement of the clutch and increased to a valuev suitable forflight on disengagement of the clutch.

43. An aircraft including an airscrew, a sustaining rotor having blades arranged so that they may be autorotated, means for imparting driving effort to said rotor, means for varying the pitch of the aircrew, means for vary n the incidence of the rotor blades, and a common control device for the first two mentioned means.

44. An aircraft including an airscrew, a sustaining rotor having blades arranged so that they may be autorotated, means for imparting driving effort to ,said rotor, means for varying the pitch of the airscrew, means for varying the incidence of the rotor blades, and a common control device for the second and third mentioned means.

45. An aircraft including an airscrew, a sustaining rotor having blades arranged so that they may be autorotated, means for imparting driving eifort to said rotor, means for varying the pitch of the airscrew, means for varying the incidence of the rotor blades, and a common control device for all threeof said means.

46. In an aircraft, a propulsion engine, an airscrew driven thereby for translational flight of the craft, a normally autorotative rotor comprising wings or blades for the sustension of the craft, a rotor starter for driving the rotor from said engine, means responsive to operation of said starter for automatically placing the rotor blades at approximately zero-lift incidence, means responsive to operation of said starter for automatically placing the airscrew at approximately zero pitch, and mechanism for automatically raising the pitch of the rotor blades and the pitch of the airscrew upon cessation of the starter drive.

47. An aircraft including a propulsion engine, a propeller normally driven thereby, a normally autorotative sustaining-bladed rotor, a disengageable transmission for starting said rotor assassa from said engine, control means for varying the pitch of the propeller, control means for operating said transmission, and mechanism controlled by operation of said transmission for controlling the rotor blade pitch angle and the propeller blade pitch in a sense such that disengagement of the transmission is accompanied by increase of rotor blade pitch angle and of propeller blade pitch angle.

48. An aircraft including a propulsion engine, a propeller normally driven thereby, a normally autorotative sustaining rotor, a disengageable transmission for starting said rotor from said engine, control means for varying the pitch of the propeller and control means for operating said transmission and means positively interrelating the operation of said two control means in a sense such that actuation of the control means to disengage the transmission is accompanied by increase of propeller blade pitch angle.

49. An aircraft including a propulsion engine, a propeller normally driven thereby, a normally autorotative sustaining rotor, a disengageable transmission for starting said rotor from said engine, control means for varying the pitch of the propeller and control means for operating said transmission and means positively interrelating the operation of said two control means in a sense such that actuation of the control means to disengage the transmission is accompanied by increase of propeller blade pitch angle, and means providing for independent control of the propeller pitch when the rotor transmission is disengaged, within a limited rangeof positive pitch variation sufficient to enable optimum propeller efliciency to be obtained in all conditions of flight,

50. An aircraft including an engine, an airscrew normally driven thereby, a sustainingbladed rotor, a disengageable transmission for driving said rotor from said engine, control means for varying the pitch of the airscrew, control means for operating said transmission, and mechanism controlled by operation of said control means for said transmission for controlling the rotor blade pitch angle and the airscrew blade pitch.

51. An aircraft including an engine, an airscrew normally driven thereby, a sustaining rotor, a disengageable transmission for driving said rotor from said engine, control means for varying the pitch of the airscrew and control means for operating said transmission,v and means positively interrelating the operation of said two control means.

52. An aircraft including an engine, an airscrew normally driven thereby, a sustaining rotor, a disengageabletransmission for driving said rotor from said engine, control means for varying the pitch of the airscrew and control means for operating said transmission, means positively interrelating the operation of said two control means, and means providing for independent control of the airscrew pitch when the rotor transmission is disengaged, within a range of positive pitch variation sufficient to enable optimum airscrew efliciency to be obtained in all conditions of flight.

53. A rotor blade articulation comprising an apertured lug member and a pivot member, one secured to the rotor hub and the other to the rotor blade, the pivot member being formed of two complementary parts fitted into the lug aperture from opposite sides, one part being ex- 

